This invention relates to a shroud assembly for a turbine engine. In particular, it relates to cooling of a shroud assembly in a gas turbine engine.
Cooling of the shroud surrounding the turbine wheel in a turbine, presents rather unique problems. The shroud surrounding the turbine blades must be in close proximity to the blades in order to maintain efficiency in the turbine engine. Notwithstanding temperature, the turbine wheel must rotate freely, both at start up and during operation. It is a characteristic of turbine engines that the turbine wheel operates at a relatively high temperature. Similarly the shrouds surrounding the turbine wheel operate at a relatively high temperature. It usually is a characteristic of turbines that the material from which the turbine wheel and turbine wheel blades are made will expand at a differing rate than the surrounding shroud structure. Although it would be possible to make the turbine wheel and blades of the same material as the shroud structure, heating of the two elements of the turbine (i.e. the wheel/blades and the shroud) will differ. Experience shows that the shroud usually operates at a higher temperature than the turbine wheel. Therefore, if the two elements are made of the same material, the shroud will expand at a faster rate and finally obtain a relatively larger inside diameter than the outside diameter of the steady state expanded turbine wheel. In this situation, some of the hot fluid powering the turbine wheel will bypass the turbine blades and further may cause unnecessary turbulence in the vicinity of the turbine blades, thus resulting in excessive fuel consumption.
Accordingly, it is desireable to provide cooling in the vicinity of the turbine shroud to decrease the difference in expansion between the turbine shroud and the turbine wheel and blades. Cooling fluid for turbine engines of the gas turbine type used in propelling aircraft is readily available either from atmospheric air flow over an uninsulated engine case, or bleed air from the compressor, or, in the case of a turbofan engine, air bled from the fan.
In an industrial gas turbine engine of the embodiment of this invention, atmospheric air or fan bleed air is not available. Furthermore, the engine case in an industrial gas turbine engine is generally heavily insulated to prevent heat loss within the engine itself, thus ambient air is of little use. Compressed air flow from the compressor stage of an industrial gas turbine engine is usually communicated directly to a heat exchanger. The heat exchanger serves two purposes, first to warm the incoming air for subsequent combustion in the gas turbine itself and secondly, to cool the exhaust gases before discharge into the atmosphere. It is impractical to utilize the compressed air from the heat exchanger with its recuperated heat for cooling of the turbine shroud since the temperature of this air is excessive. On the other hand, air may be bled directly from the compressor stage and communicated through appropriate manifolding to cool the various gasifier turbine parts. This compressor bleed air has a relatively cool temperature established primarily by the compression ratio, and secondarily by conduction from the hot engine case.
Use of bleed air from the turbine compressor should be limited in order to achieve the highest degree of engine efficiency. In earlier industrial gas turbine engines, air was supplied in a random manner to the shroud structure surrounding the turbine blades. Furthermore, the material utilized in earlier shroud structures was usually chosen primarily for its strength with secondary attention paid the coefficient of thermal expansion.
Furthermore, in earlier industrial gas turbine engines wherein the cooling air was provided in an haphazard fashion to the shroud rings, no attempt was made to isolate the cooling air from the surrounding hot structure thus; by the time the cooling air arrived at the shroud structure, a good deal of its cooling potential had been lost due to temperature increases through contact with hot surfaces of the gas turbine engine. Finally, the large plenum chamber arrangement involved in earlier gas turbine engines resulted in a drop in pressure of the cooling air to the extent that hot gas was able to enter the plenum chamber and further degrade the cooling.
Attempts to maintain a smooth gas flow through the turbine and past the turbine wheel, resulted in the turbine shroud essentially being made an integral part of the turbine casing. Thus, the adjacent relatively high temperatures of the turbine casing were conducted to the turbine shroud with a concomitant expansion of the turbine shroud. Even though attempts have been made to restrain the expansion of the turbine shroud, efforts along this line have not been entirely successful. To compound the problem, reduction in a diameter or maintenance of the same diameter of the turbine shroud ring resulting from impingement of cooling air thereon was resisted by the mechanical restraints imposed on the turbine shroud ring by expansion of the hotter portions of the turbine engine casing.
Finally, earlier gas turbine engines, both of the industrial type and the aircraft type, usually used an overlapping segmented shroud assembly to permit thermal expansion within each segment while not substantially increasing the inside diameter of the entire shroud structure. These individual shroud segments were mounted in various ways with cooling air usually directed toward them in the aforedescribed haphazard manner. A possible disadvantage resulted from such a structure. A true circular opening for the turbine wheel was difficult to achieve because of the segmented nature of the shroud assemblies themselves. Therefore, the clearance between the turbine shroud assembly and a turbine blade had to be adjusted to account for a possible out of round condition. This adjustment resulted in an unnecessary loss of efficiency.